low-Modulus Gas-Turbine Compressor Blade

ABSTRACT

The present invention relates to an aircraft gas-turbine compressor blade having an airfoil with a leading edge, with at least the area of the leading edge being made from a low-modulus titanium alloy.

This invention relates to a gas-turbine compressor blade, in particularto an aircraft gas turbine or a stationary turbine, having an airfoilfastened to a blade root. The compressor blade has a leading edge orinflow edge, which is also called the blade front edge.

Aircraft gas turbines always face the problem that the compressor bladesare subjected to heavy erosion due to sucked-in particles. These are forexample grains of sand or the like which impact the compressor blades athigh velocity.

As a result of the erosion occurring, it is necessary to replace orrepair the compressor blades. Repair is, in the case of blisks inparticular, very complex from the engineering viewpoint and also verycost-intensive.

The erosion-resistant materials known from the state of the art areunsuitable as materials for compressor blades. Erosion-resistantcoatings too have proven in practice to be unsuitable, as they are onlyeffective for very small particle sizes.

The object underlying the present invention is to provide an aircraftgas-turbine compressor blade which, while being simply designed andeasily and cost-effectively producible, has a high erosion resistance.

It is a particular object of the present invention to provide solutionto the above problematics by a combination of the features of Claim 1.Further advantageous embodiments of the invention become apparent fromthe sub-claims.

It is provided in accordance with the invention that at least the areaof the leading edge of the compressor blade is made from a materialwhich has a very low modulus of elasticity. A suitable and low modulusof elasticity is in the range of 50 GPa.

The compressor blade in accordance with the invention thus has, due tothe low modulus of elasticity, a relatively high elasticity in the areaof the leading edge, so that an impact of a foreign object, for examplea sand particle or stone, does not lead to damage thanks to theelasticity of the material in the area of the leading edge. Inparticular, zero or only very low erosion occurs when compared withcompressor blades known from the state of the art.

In accordance with the invention, it is particularly advantageous whenat least the area of the leading edge is made from a titanium alloy oftypes Ti-28Nb-1Fe-0.5Si, Ti-36Nb-2Ta-3Zr-0.3O or Ti-24Nb-4Zr-8Sn. Bymeans of an alloy of this type, it is for example possible to reduceerosive wear by about 50% compared with compressor blades made from aconventional alloy, for example Ti64. Compared with compressor bladesknown from the state of the art made from nickel-based alloys, forexample IN718, erosive wear can be reduced by 30%.

It is possible in accordance with the invention to either manufacturethe entire compressor blade from the above titanium alloy with the lowmodulus of elasticity or only to manufacture parts of the compressorblade from this material.

In a particularly favourable embodiment of the invention, it is providedthat the leading edge is designed in the form of a leading-edge elementwhich is connected to the airfoil of the compressor blade by means of awelding process. Particularly advantageous is the use of a laser weldingprocess.

The airfoil itself can, in accordance with the invention, be made fromone of the conventional titanium alloys, for example Ti64, Ti6246 orTi6242.

If a separate leading-edge element is provided, the latter extendspreferably up to just in front of the blade root, and in the directtransition of the blade root it is not necessary to apply or use thealloy in accordance with the invention.

In the case of separate manufacture of the leading-edge element inaccordance with the invention and connection by means of a laser weldingprocess, there is the advantage that due to the thermal influence duringthe laser welding operation a soft transition is obtained between thelow modulus of elasticity of the leading-edge element and the highermodulus of elasticity of the material in the rest of the airfoil. Theresult is thus a gradual transition of the elastic properties, whichacts advantageously on the behaviour of the entire compressor bladecomponent. This effect is based on the fact that during heating of thelow-modulus titanium alloy in temperature ranges above around 500° C.,the modulus of elasticity rises again to values of conventional titaniumalloys, and hence the direct area of the weld is homogeneous.

The compressor blade in accordance with the invention, which can bedesigned as a single compressor blade, as a blisk blade or as a fanblade, is thus to a high degree resistant against erosion and damagefrom foreign objects.

The invention thus also relates to a method for the manufacture of acompressor blade in which a leading-edge element is manufacturedseparately and is connected to the airfoil by means of a weldingprocess, in particular by a laser welding process, where theleading-edge element is made from a low-modulus (in respect of themodulus of elasticity) titanium alloy, for example fromTi-28Nb-1Fe-0.5Si, from Ti-36Nb-2Ta-3Zr-0.3O or from Ti-24Nb-4Zr-8Snand/or an alloy with a modulus of elasticity of substantially 50 to 70GPa and where the airfoil is made from an alloy known from the state ofthe art.

The invention also relates to the use of a low-modulus (of elasticity)titanium alloy, for example Ti-28Nb-1Fe-0.5Si or Ti-36Nb-2Ta-3Zr-0.3O orTi-24Nb-4Zr-8Sn, in particular for the leading-edge area of a compressorblade.

The present invention is described in the following in light of theaccompanying drawing, showing an exemplary embodiment. In the drawing,

FIG. 1 shows a schematized representation of a gas-turbine engine inaccordance with the present invention,

FIG. 2 shows a schematized side view of a compressor blade in accordancewith the present invention, and

FIG. 3 shows a representation of the curve of the modulus of elasticityas a function of the blade length.

The gas-turbine engine 10 in accordance with FIG. 1 is an example of aturbomachine where the invention can be used. The following howevermakes clear that the invention can also be used in other turbomachines.The engine 10 is of conventional design and includes in the flowdirection, one behind the other, an air inlet 11, a fan 12 rotatinginside a casing, an intermediate-pressure compressor 13, a high-pressurecompressor 14, combustion chambers 15, a high-pressure turbine 16, anintermediate-pressure turbine 17 and a low-pressure turbine 18 as wellas an exhaust nozzle 19, all of which being arranged about a centralengine axis 1.

The intermediate-pressure compressor 13 and the high-pressure compressor14 each include several stages, of which each has an arrangementextending in the circumferential direction of fixed and stationary guidevanes 20, generally referred to as stator vanes and projecting radiallyinwards from the engine casing 21 in an annular flow duct through thecompressors 13, 14. The compressors furthermore have an arrangement ofcompressor rotor blades 22 which project radially outwards from arotatable drum or disk 26 linked to hubs 27 of the high-pressure turbine16 or the intermediate-pressure turbine 17, respectively.

The turbine sections 16, 17, 18 have similar stages, including anarrangement of fixed stator vanes 23 projecting radially inwards fromthe casing 21 into the annular flow duct through the turbines 16, 17,18, and a subsequent arrangement of turbine blades 24 projectingoutwards from a rotatable hub 27. The compressor drum or compressor disk26 and the blades 22 arranged thereon, as well as the turbine rotor hub27 and the turbine rotor blades 24 arranged thereon rotate about theengine axis 1 during operation.

FIG. 2 shows a schematic side view of a compressor blade in accordancewith the invention, which can be a conventional compressor blade, a fanblade or a blisk blade. The blade has for example a blade root 31 towhich is connected an airfoil 29 designed in the usual way. Aleading-edge element 30 is connected by means of a laser weld 32 in thearea of a leading edge or inflow edge. The leading-edge element 30extends up to just in front of the blade root 31, and does not need tobe connected directly to the blade root 31.

FIG. 3 shows a representation of the curve of the modulus of elasticityin schematic form. It can be seen that the modulus of elasticity of theleading-edge element 30 is lower than the modulus of elasticity of theairfoil 29, where a soft transition is obtained in the area of the weld32 between the lower and the higher moduli of elasticity.

LIST OF REFERENCE NUMERALS

-   1 Engine axis-   10 Gas-turbine engine/core engine-   11 Air inlet-   12 Fan-   13 Intermediate-pressure compressor (compressor)-   14 High-pressure compressor-   15 Combustion chambers-   16 High-pressure turbine-   17 Intermediate-pressure turbine-   18 Low-pressure turbine-   19 Exhaust nozzle-   20 Guide vanes-   21 Engine casing-   22 Compressor rotor blades-   23 Stator vanes-   24 Turbine blades-   26 Compressor drum or disk-   27 Turbine rotor hub-   28 Exhaust cone-   29 Airfoil-   30 Leading-edge element-   31 Blade root-   32 Weld-   33 Leading edge

1. Gas-turbine compressor blade having an airfoil with a leading edge,with at least the area of the leading edge being made from a low-modulustitanium alloy.
 2. Compressor blade in accordance with claim 1,characterized in that at least the area of the leading edge is made fromTi-28Nb-1Fe-0.5Si or Ti-36Nb-2Ta-3Zr-0.3O or Ti-24Nb-4Zr-8Sn. 3.Compressor blade in accordance with claim 1, characterized in that theleading edge is designed in the form of a leading-edge element which isconnected to the airfoil by means of a welding process.
 4. Compressorblade in accordance with claim 3, characterized in that the leading-edgeelement is connected to the airfoil by means of a laser welding process.5. Compressor blade in accordance with claim 1, characterized in thatthe area of the leading edge has a modulus of elasticity ofsubstantially 50 to 70 GPa.
 6. Compressor blade in accordance with claim1, characterized in that the airfoil is made from Ti64 or Ti6246 orTi6242 or IN718.
 7. Compressor blade in accordance with claim 1,characterized in that the compressor blade is designed as a conventionalcompressor blade.
 8. Compressor blade in accordance with claim 1,characterized in that the compressor blade is designed as a fan blade.9. Compressor blade in accordance with claim 1, characterized in thatthe compressor blade is designed as a blisk blade.